Starting of a small turbojet

ABSTRACT

Difficulties in starting a turbine engine, particularly ones having a small volume combustor (83) with a low dome height, may be avoided by a method including the steps of a) forming the turbine engine compressor (48) of a low density, high temperature resistant ceramic material, b) directing pyrotechnic gases at a temperature of at least about 3,000° F. against the blades (48) of the compressor, c) continuing the step of directing until the turbine rotor (44) is rotating at at least about 60 percent of rated speed and d) combusting fuel in the combustor (83) of the turbine engine to maintain the rotor (44) rotating at at least about 60 percent of rated speed.

This is a division of application Ser. No. 07/612,109 filed Nov. 9,1990, U.S. Pat. No. 5,263,315.

FIELD OF THE INVENTION

This invention relates to a small, trust producing, gas turbine engineand more particularly, to the starting thereof.

BACKGROUND OF THE INVENTION

Most relatively small missiles in use today are propelled by solid fuelrockets as opposed to, for example, turbojet engines. The selection of asolid fuel rocket as a propulsion device has been largely dictated bytwo factors. First, in many instances, a turbine engine cannot befabricated sufficiently economically as to compete with a solid fuelrocket engine. Secondly, in small size missiles, i.e., those having arelatively small diameter measured on the order of about six inches orless, it has heretofore been extremely difficult to manufacture anefficient turbojet engine that will fit within the envelope required forthe propulsion unit for such a missile.

As a consequence of the resulting use of solid fuel rocket engines, somedegree of control of the missile flight path or trajectory is lost overthat which is available were it possible to propel the missile by a gasturbine engine whose output can be readily varied. Further, even if thegas turbine engine operates relatively inefficiently, the use of such anengine greatly extends the range of the missile.

Recently, in order to overcome the difficulties attendant the use ofrocket engines, the assignee of the instant application has produced agas turbine engine having a diameter of about six inches. These enginesare disclosed in U.S. Letters Pat. No. 4,794,754 issued Jan. 3, 1989 toShekleton, et al and have been extremely successful in meeting thechallenge of providing a propulsion unit for small diameter missiles.

To achieve low cost, in order to replace solid rockets as propulsionsources for such missiles, it is necessary that not only the turbojetengines themselves be manufactured relatively inexpensively, but theirperipheral systems must likewise be manufactured relativelyinexpensively. At the same time, a high degree of reliability isrequired. The system employed for starting such a turbojet engine is noexception. Thus, the present invention is directed to fulfilling theneed for a highly reliable, low cost starting system for a thrustproducing gas turbine engine, particularly those that are low cost andexpendable.

SUMMARY OF THE INVENTION

It is the principal object of the invention to provide a new andimproved gas turbine engine with a low cost starting system. It is alsoan object of the invention to provide a new and improved method ofstarting a gas turbine engine.

The invention achieves one or more of the above objects, according toone facet thereof, in a gas turbine engine which includes a centrifugalcompressor having blades formed of a material having a low density and ahigh temperature resistance. A turbine wheel is coupled to thecompressor to drive the same and a nozzle is provided for directing thegases of combustion at the turbine wheel during normal operation. Acombustor is provided for receiving compressed air from the compressorand fuel from a fuel source to generate the gases of combustion. Forstarting purposes, a pyrotechnic cartridge for generating hot gases isprovided along with at least one start nozzle which is directed towardthe compressor blades. Means connect the cartridge and the start nozzleso that hot gases may be directed at the compressor blades to acceleratethe same to start the engine. The high temperature resistance of thematerial of which the compressor blades are formed allows the hot gasesto impinge on the blades without damage thereto to achieve rapidacceleration while the low density of the material likewise allows rapidacceleration of the compressor and the turbine wheel coupled thereto toa high percentage of rated speed to achieve a stable combustion point.

The rapid acceleration of the turbine wheel that results from theforegoing also assures that the blades will not be subject to the hotgases for a prolonged period of time to thereby further minimize thepossibility of damage thereto.

In one embodiment of the invention, the turbine wheel is alsoconstructed of low density material and in a highly preferred embodimentof the invention, the low density material is a ceramic.

The invention contemplates that the blades have radially outer dischargeends and that one or more start nozzles be directed at the dischargeends. As a consequence of this construction, the acceleration forces ofthe hot gases are acting over a longer moment arm than would be the caseif the hot gases were directed at the compressor blade well inwardly oftheir radially outermost extent.

In one embodiment, the start nozzle includes a converging sectioncoupled to the connecting mean and a diverging section adjacent thecompressor blades. The two are interconnected by a sonic throat. As aconsequence of this construction, the gases impinging against thecompressor blades will be at supersonic velocity.

In a highly preferred embodiment, to provide for radial compactness ofthe turbine engine, the compressor is a mixed flow compressor and thenozzle is axially directed.

In a highly preferred embodiment, a ceramic monorotor defines both thecentrifuged compressor and the turbine wheel.

According to another facet of the invention, there is provided a methodof starting a turbine engine having a combustor which includes the stepsof forming the turbine engine compressor of a low density, hightemperature resistant ceramic material, directing a high temperaturepyrotechnic gas against the blades of the compressor, continuing thestep of directing the gas until the turbine rotor is rotating at apredetermined percent of rated speed and combustion fuel in thecombustor of the turbine engine to maintain operation thereof.

In a preferred embodiment of the invention, the pyrotechnic gases are atleast at about 3,000° F. and the step of directing the same against theblades of the compressor is maintained until the turbine rotor isrotating at at least about 60 percent of rated speed.

In one embodiment of the invention, the step of combusting fuel isinitiated by igniting fuel in the combustor by the pyrotechnic gases.

The invention contemplates the step of directing pyrotechnic gases athigh temperature against the blades of the compressor also include thestep of directing the pyrotechnic gases against the blades of thecompressor at a supersonic velocity.

Other objects and advantages will become apparent from the followingspecification taken in connection with the accompanying drawings.

DESCRIPTION OF THE DRAWINGS

FIG. 1 is a sectional view of a gas turbine engine embodying a startingsystem made according to the invention; and

FIG. 2 is an enlarged, fragmentary sectional view of a starting systememployed in the engine; and

FIG. 3 is a typical plot of the inability to maintain stable andefficient combustion within a turbine engine versus percent of ratedengine speed.

DESCRIPTION OF THE PREFERRED EMBODIMENT

An exemplary embodiment of a small diameter gas turbine, thrustproducing engine is illustrated in FIG. 1. Generally, the engine will bein a form of a thrust producing engine, but it should be understood thatin some instances, the engine may be utilized in an environment wherepower is taken from a rotation shaft coupled to the rotor of the engine.

The gas turbine engine of the invention includes a bell shaped, airinlet housing 10 having a forwardly facing open end 12 through which airmay enter the engine. A plurality of struts 14 (only one of which isshown) extend radially inwardly to support a bearing and accessoryhousing 16 centrally within the open end 12. The housing 16 includesaxially spaced bearing 18 and 20 which journal a shaft 22 for rotationabout an axis 24. One end 26 of the shaft is coupled to a fuel pump 28for the engine which is adapted to receive fuel on an inlet line 30 froma tank (not shown) and pump the fuel through an outlet 32 to a fuelcontrol system (not shown) which in turn provides fuel for combustionwithin the engine to a fuel manifold 34 to be described in greaterdetail hereinafter.

Also contained within the housing 16 is a small permanent magnetgenerator, generally designated 36, which is utilized to generateelectrical power for operating engine and missile controls (not shown).

The end 38 of the shaft 22 opposite the end 26 includes a recess 40which receives a protuberance 42 on one end of an engine rotor 44. Therotor 44 is a so-called monorotor in that it is of one piececonstruction. The same is cast of a low density material, preferably ofceramic material so that the rotor will have a low mass and thus may beaccelerated rather quickly during a starting sequence. Ceramic materialalso has good resistance to high temperatures thus allowing a higheroperation temperatures for more rapid, reliable starts as will be seen.Preferably, the rotor 44 is brazed to the shaft 22 at the interface ofthe recess 40 and the protuberance 42.

An end 46 of the rotor 44 adjacent the protuberance 42 mounts aplurality of integral compressor blades 48 which thus are formed ofceramic. The blades 48 include inlet edges 50 and discharge ends 52. Ascan be seen from FIG. 1, the blades 48 are configured to provide a mixedflow rotary compressor which is to say that at the discharge end 52, theflow of compressed gas has appreciable or substantial radial componentsas well as axial components. This is in contrast to conventional gasturbine engines wherein compressor discharge is either substantiallyaxial or substantially radial.

The end 54 of the rotor 44 opposite the end 46 mounts a plurality ofintegral turbine blades 56 which likewise are formed of ceramic. Theturbine blades 56 have inlet edges 58 which are generally parallel tothe axis 24 and downstream edges 60 which typically, but not always,will be generally transverse to the axis 24. Those skilled in the artwill immediately appreciate that the blades 56 having the configurationdescribed define a radial turbine wheel.

Hot gases of combustion directed against the blades 56 are confined tothe spaces between adjacent blades 56 by a rear shroud 62 which, as canbe seen in FIG. 1, is axially elongated. Spaced radially outward of theshroud 62 is a sleeve-like partition 64 which is axially elongated andhas one end 66 in overlying relation to a land 70 on the rotor 44separating the compressor blades 48 from the turbine wheel blades 56. Asbest seen in FIG. 2, a series of radially inwardly opening, annulargrooves 72 are located in the radially inner side of the ends 66 of thepartition 64 and are separated by annular points or projections 74 thatalmost, but do not quite, contact the land 70. Thus, a labyrinth seal isdefined at the interface of the partition end 66 and the land 70 and iscarried by the partition 64.

The end 68 of the partition 64 extends axially the full length of theshroud 62 and a series of axially elongated turbine nozzle blades 76 aresupported between the shroud 62 and the partition 64 to define an axialnozzle, generally designated 77. The axially elongated nozzle blades 76direct hot gases of combustion axially through the space between theshroud 62 and the partition 64 against the blades 58 of the turbinewheel defined thereby to drive the same. In this connection, it will beobserved that an annular, concave surface 78 is located intermediate theends of the partition 64 to direct the axially flowing gases from theblades 76 radially against the turbine blades 56 at their inlet edges58.

The use of axially elongated blades 76 having leading edges 80 axiallyspaced from the trailing edges 82 in contrast to the conventionalarrangement for nozzle blades in a radial turbine construction allowsreduction of the engine diameter. In the conventional case, the trailingedges of the vanes would be radially inwardly of the leading edges andthe outlet of a combustor would be radially outwardly of the leadingedges. This adds sizable to the diameter of the machine in contrast tothe configuration illustrated wherein a reverse flow, annular combustor,generally designated 83, includes a radially inner wall joined to theshroud 62 and a radially outer wall 86 joined to the partition 64. Thespace between the radially inner and outer walls 84, 86 at the shroud 62and the partition 64 define the outlet of the combustor 83 and it willbe observed that such outlet has essentially the same radius as theturbine nozzle defined by the blades 76. This clearly illustrates howdiametral compactness is achieved through the use of the axial blades76.

A cylindrical engine case 89 surrounds the combustor 83 and is spacedradially outwardly therefrom. In the vicinity of the partition 64, thesame supports, together with the partition 64, a two row, axial, cascadediffuser, generally designated 90, forming no part of the invention. Thediffuser 90 includes rows of diffuser vanes 90 and 94 in close proximityto the discharge ends 54 of the compressor blades 48. The configurationof the vanes 92 and 94 may be achieved through conventional art and itwill be observed that in the downstream row of vanes, there are twice asmany vanes as there are in the first row. Such a configuration extendsthe stall range.

The case 89 extends about the radially outer wall 86 of the combustor 83in spaced relation to define an annular plenum 98 which receives thedischarge from the diffuser 90. As can be seen in FIG. 1, the combustor83 is axially elongated and relatively narrow. A radially extending wall100 interconnecting the radially inner and outer walls 84, 86 remotefrom the turbine nozzle blades 76 defines the so-called "dome" of thecombustor 83 and it will be appreciated that the so-called dome heightis but a minor fraction of the axial length of the combustor 83. In theillustrated embodiment, (which is approximately to scale in FIG. 1) thedome height is about 1/3 of the axial length of the combustor 83.

Three rows of air injection tubes 102, 104, 106 are axially spaced fromone another and arranged to establish fluid communication through theradially outer wall 86 between the plenum 98 and the interior of theannular combustor 83. In a typical case, there will be perhaps four ofthe tubes 102 in a first row nearest the dome or radially extending wall100, an equal number of the tubes 106 adjacent the nozzle 77 and acommensurate number of intermediate tubes 104 located between the rowsof tubes 102 on the one hand and the rows of tubes 106 on the other.

In the usual case, 1/3 of the air introduced into the combustor 83 willenter through the tubes 106, another 1/3 through the tubes 104 and the1/3 through the tubes 102. The invention takes advantage of this featureof the combustor 83 to maximize the combustor volume. In particular, theradially outer wall 86 is somewhat frusto-conical, diverging outwardlytoward the case 89 as one progresses toward the dome or radiallyextending wall 100. Thus, the plenum 83 progressively is narrowed, butsuch is entirely permissible since there is a considerably lesser volumeof air flowing in the plenum 83 adjacent the row of tubes 102 than thereis flowing adjacent the row of tubes 106. As is well known, theincreased combustor volume allows the achievement of a greater powerlevel without loss of stability in flame propagation.

The tube 102 has an axis 110 which is transverse to the rotational axis24 as can be ascertained from FIG. 1 and which is tangential to thespace between the radially inner and outer wall 84, 86. The axes of thetubes 104 and 106 are similarly oriented. Thus, the air introduced intothe combustor 83 via the tubes 102, 104 and 106 will be movingcircumferentially to create a swirling flame. As has been recognized inrecent years, the use of swirl within the combustion chamber providesfor a stable flame while minimizing the number of fuel injectors andprovides the residence time needed to complete combustion in small sizecombustors by lengthening the path of the fuel and air mixture ascombustion takes place.

Fuel injection manifolds 111 associated with the tubes 102 and 104(FIG. 1) and connected to the manifold 34 may have apertures associatedwith each of the tubes 102 and 104 within the plenum 98 as more fullydisclosed in the previously identified Shekleton, et al. patent. As aresult, fuel may be discharged into the air stream entering such tubesto be entrained therewith and directed along with air into the combustor83.

The radially inner wall 84 of the combustor 83 may be extended past theradially extending wall 100. Such an extension is designated 120 andserves as an exhaust duct for the engine. The wall 100 may likewisesupport a conventional ignitor 122 which will extend into the interiorof the combustor 83 to achieve ignition therein.

For starting purposes, a gas discharge port 126 may be located in acontinuation 128 of the inlet housing 10 which additionally serves as ashroud for the compressor blades 48. A conduit 130 may be connected to asource of high pressure gas or the like and when such gas is allowed toflow through the opening 126, it will impinge against the blades 48 torapidly accelerate the rotor 44 to a speed whereat ignition may be hadand stably maintained.

More specifically, according to the invention, the source of highpressure gas is in fact a conventional pyrotechnic cartridge 132 which,when ignited, will provide extremely high pressure gases at hightemperatures of at least about 3,000° F. and conceivably as high as5,000° F. at the compressor blades 48. These gases are passed to the gasdischarge port 126 which acts as a nozzle for focusing the hotpyrotechnic gases against the compressor blades 48 near their dischargeedges 52.

In fact, there may; be a plurality of the gas discharge ports 126,preferably at least three, equally angularly spaced about the axis 24(FIG. 1). Each includes a converging section 134 connected to thepyrotechnic cartridge 132 by the conduit 130 and a diverging section 136closely adjacent to the discharge ends 52 of the blades 48. A sonicthroat 138 interconnects the converging end 134 with the diverging end136.

As a result of the foregoing, not only will the hot gases from thepyrotechnic cartridge be applied to the blades 48 at a high temperatureof at least 3,000° F., after passing through the throat 138 at sonicvelocity, they will have expanded rapidly within the diverging section136 and achieved supersonic velocity at the time of application to theblades 38 for maximum efficiency and rapid acceleration.

It will be observed that the nozzles defined by the gas discharge ports126 are axial and act against the radially outer ends of the blades 48.This maximizes the moment arm from the point of application of the hotgases to the rotational axis of the rotor to assure extremely rapidacceleration to minimize the time that each individual blade 48 isexposed to hot gases coming out of the ports 126. It also provides ameasure of convenience in manufacture of the engine, particularly whenthe same is an extremely small diameter engine of, say, for example,four inches in diameter.

In conventional turbines of this sort, the maximum temperature of gasesdirected against rotor blades such as the turbine blades 56 are atrelatively low temperatures of about 2,100° F. At higher temperatures,thermal damage to the metal of which the blades are formed may occur.When metals of greater temperature resistance are utilized to allowhigher gas temperatures, the acceleration rate of the rotor duringstarting is generally slowed since the metals that are more resistant tohigher temperatures typically have higher densities. With the resultingslowing of acceleration, the blades are exposed to the hot gases for alonger time and that in turn further inhibits the use of lesstemperature resistant material or those having relatively highdensities.

According to the invention, by forming at least the blades 48 of thecompressor end 46 of the rotor 44 of ceramic, and preferably, formingthe turbine blades 56 of the turbine end 54 of the rotor 44 of ceramicas well, and even more preferably, in forming the entire rotor 44 as amonorotor with integral blades 48 and 56 of ceramic, the density of therotating components of the engine may be minimized. This in turnincreases the rate of acceleration of the rotatable components of theengine which in turn reduces the length of exposure of the components tothe high temperature gases allowing an increase in the temperature ofthe gases applied thereto. And coupled with a higher temperatureresistance provided by a ceramic, it will be appreciated that quite hightemperatures may be utilized to achieve rapid acceleration of the rotor44 well into a stable operating range in short periods of time that maybe as little as one second.

FIG. 3 illustrates a typical plot of the inability to maintain stableand efficient combustion within a gas turbine engine versus a percent ofrated engine speed. In typical turbines, the greatest difficulty (worstcase) in maintaining efficient and stable combustion occurs atintermediate speeds of approximately 40 percent of rated engine speed.If the turbine engine is a so-called low pressure ratio engine, theworst case would occur at a somewhat higher speed whereas if the engineis a so-called high pressure ratio engine, the worst case will occur ata somewhat lower speed. In any event, a typical worst case condition isillustrated at a point 140.

Because of the rapid acceleration achieved by the invention, it is nowpossible to accelerate the rotor up to 60 or even 80 percent of ratedspeed before combustion is initiated. This is shown as point 142 in FIG.3 and it will be immediately recognized by those skilled in the art thatthe engine is almost immediately operating in an efficient and stableoperating zone. So long as engine speed is maintained at 60 percent ofrated speed or more, highly stable operation is achieved.

Another advantage may accrue as a result of the high temperature of thegases employed to start the engine. It will be appreciated that suchgases, after impinging against the discharge ends 52 of the compressorblades 48 ultimately follow the route of incoming air and pass throughthe discharge duct 120 of the engine which, of course, means that theyhave passed through the combustor 83 in the process. In many instances,the temperature of the hot gases at this point in time will besufficient to cause ignition of any fuel admitted to the interior of thecombustor 83 so that the ignitor 122 may be eliminated, thus reducingcomplexity as well as cost.

Still another advantage may occur. Because the hot gases are flowingthrough the engine in the manner mentioned immediately preceding, theyimmediately warm up the parts thereof, particularly those withrelatively low mass such as the walls 84, 86 and 100 of the combustor83. Thus, such walls are at an elevated temperature at the timecombustion is initiated and that in turn reduces the difficulties ofachieving stable combustion in extremely small combustors as may beemployed when the engine is on the order of four inches in diameter.

We claim:
 1. A method of starting a turbine engine having a rotor, acombustor for supplying hot gases to cause rotation of said rotor, arotatable compressor for supplying gases to said combustor, saidcompressor having a plurality of blades extending outwardly therefrom, afuel source for supplying fuel to said combustor, comprising:a) prior tooperation of said turbine engine, forming at least said blades of thecompressor of ceramic material; b) thereafter, directing pyrotechnicgases having a temperature of at least about 3,000° F. and a supersonicvelocity against said blades of the compressor until said rotor hasreached a predetermined rotational speed; c) introducing fuel into saidcombustor and igniting that fuel to cause combustion to produce said hotgases; and d) continuing said combustion to maintain said predeterminedrotational speed of said rotor after said directing of pyrotechnic gasesceases.
 2. The method according to claim 1 wherein said pyrotechnicgases are also directed into said combustor to cause said igniting offuel.
 3. A method of starting a turbine engine having a rotor, acombustor means for supplying hot gases to cause rotation of saidturbine rotor, a rotatable compressor means for supplying gases to saidcombustor means and a fuel supply means for providing fuel to saidcombustor means for combustion therein upon mixing with said gases fromsaid compressor means, comprising the steps of:a) initially directingpyrotechnic gases at said compressor means so as to cause rotation ofsaid compressor means; b) thereafter, directing said pyrotechnic gasesinto said combustor means so as to cause ignition of said fuel thereinto begin combustion to supply hot gases to said turbine rotor sufficientto cause rotation thereof; and c) continuing to direct said pyrotechnicgases at said compressor means until said turbine rotor is rotating at apredetermined speed, that speed being sufficient to maintain stablecombustion within said combustor without use of said pyrotechnic gases.4. The method of claim 3 wherein said pyrotechnic gases are directedagainst the compressor means at a supersonic velocity.
 5. The method ofclaim 3 wherein said compressor means and said turbine rotor areintegrally formed as a monorotor structure.